This invention generally relates to gas turbine engines and more specifically to a shroud section that surrounds a stage of rotating airfoils in the turbine of a gas turbine engine.
A gas turbine engine typically comprises a multi-stage compressor, which compresses air drawn into the engine to a higher pressure and temperature. A majority of this air passes to the combustors, which mix the compressed heated air with fuel and contain the resulting reaction that generates the hot combustion gases. These gases then pass through a multi-stage turbine, which drives the compressor, before exiting the engine. A portion of the compressed air from the compressor bypasses the combustors and is used to cool the turbine blades and vanes that are continuously exposed to the hot gases of the combustors. In land-based gas turbines, the turbine is also coupled to a generator for generating electricity.
In the turbine section of the engine, alternating stages of rotating and stationary airfoils are present through which the hot combustion gases expand as they turn the rotating stages of the turbine. In order to maximize the performance of the turbine, it is critical to maximize the amount of hot combustion gases passing through the airfoils, and not leaking around the airfoils, nor being used to cool the airfoils. To prevent leakage around stages of rotating airfoils, or turbine blades, shroud segments are used that conform to the radial profile of the turbine stage and are sized such that when the blade is rotating and at its operating temperature, the gap between the turbine blade tip and the shroud segment is minimized.
Given that operating temperatures within the turbine typically exceed 2000 degrees F. it is necessary to provide a source of cooling to the blades, vanes, and shroud segments adjacent the rotating blades so that these components are maintained within their material operating limits. Of particular concern with respect to the present invention is cooling of the shroud segments that encompass the rotating turbine blades. However, while it is necessary to cool the shroud segments, any air directed to cool the shroud segments does not pass through the turbine, thereby reducing the turbine efficiency. It is imperative that this cooling air, which is typically drawn from the engine compressor, be a minimal amount and used most effectively to cool as much of the exposed shroud surface as possible. An example of a shroud segment for a gas turbine engine employing a form of cooling of the prior art is shown in perspective view in FIG. 1. Shroud 10 includes an inner surface 11 that faces directly towards the tips of the rotating turbine blades (not shown) and an outer surface 12 in spaced relation to inner surface 11. Extending axially through the shroud thickness between inner surface 11 and outer surface 12 and exiting from shroud aft face 13 is a plurality of cooling holes 14. A cooling fluid, such as compressed air, enters cooling holes 14 from air inlets 15 and cools the shroud 10 as it passes through cooling holes 14. In this configuration, the edges 16 and 17 of shroud 10 do not receive any dedicated cooling. Shrouds are typically segmented, creating edges 16 and 17, in order to allow for differing thermal expansion between shroud 10 and the engine case in which the shrouds are mounted. Inspection of prior art shrouds having this cooling configuration indicate excessive heat load along edges 16 and 17, especially along the axial region of shroud 10 where the turbine blade is located.
In order to overcome the shortfalls of the prior art shroud design, it is necessary to provide a shroud for a gas turbine engine which addresses the heat load issues found in the prior art design, including providing sufficient cooling to the edges of the turbine shroud. Providing sufficient cooling to the edge regions where it is most needed will ensure that the heat load is reduced in the effected areas thereby extending the life of turbine shroud segments.